This invention relates to a gas turbine airfoil having an internal cooling passage, and, more particularly, to the selective protection of the surface of the internal passage of such a gas turbine airfoil.
In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine by impingement against an airfoil section of the turbine blades and vanes, which turns the shaft and provides power to the compressor and fan. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forwardly.
The hotter the combustion and exhaust gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the combustion and exhaust gas temperatures. The maximum temperature of the combustion gases is normally limited by the materials used to fabricate the turbine vanes and turbine blades of the turbine, upon which the hot combustion gases impinge. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at temperatures of up to about 1900-2100xc2x0 F.
Many approaches have been used to increase the operating temperature limit of the turbine blades and vanes to their current levels. For example, the composition and processing of the base materials themselves have been improved.
Physical cooling techniques may also be used. In one technique, internal cooling passages through the interior of the turbine airfoil are present. Air is forced through the cooling passages and out openings at the external surface of the airfoil, removing heat from the interior of the airfoil and, in some cases, providing a boundary layer of cooler air at the surface of the airfoil. To attain maximum cooling efficiency, the cooling passages are placed as closely to the external surface of the airfoil as is consistent with maintaining the required mechanical properties of the airfoil, to as little as about 0.020 inch in some cases.
In another approach, a protective layer or a ceramic/metal thermal barrier coating (TBC) system is applied to the airfoil, which acts as a substrate. The protective layer with no overlying ceramic layer (in which case the protective layer is termed an xe2x80x9cenvironmental coatingxe2x80x9d) is useful in intermediate-temperature applications. The currently known protective layers include diffusion aluminides and MCrAlX overlays. A ceramic thermal barrier coating layer may be applied overlying the protective layer on the external airfoil surface, to form a thermal barrier coating system (in which case the protective layer is termed a xe2x80x9cbond coatxe2x80x9d). The thermal barrier coating system is useful in higher-temperature applications. The ceramic thermal barrier coating insulates the component from the combustion gas, permitting the combustion gas to be hotter than would otherwise be possible with the particular material and fabrication process of the substrate.
The surfaces of the internal cooling passages may be protected with a diffusion aluminide coating, which oxidizes to an aluminum oxide protective scale that inhibits further oxidation of the internal surfaces. Although techniques are known for depositing an aluminide protective coating on an internal passage, the present inventors have observed that the available techniques suffer from the shortcoming that they may adversely affect the protection and the repair of the external surface of the airfoil. There is a need for an improved approach to the protection of the internal cooling passages of gas turbine airfoils, which approach does not adversely affect other portions of the airfoils. The present invention fulfills this need, and further provides related advantages.
The present invention provides an apparatus and method for selectively coating the internal cooling passages of an airfoil section, while not coating the external surface with any substantial amount of the material used to coat the internal cooling passages. In the preferred practice, the surfaces of the internal cooling passages are coated with a diffusion aluminide (a term which includes composition-modified aluminides as used herein), which is later oxidized to form a protective aluminum oxide scale. Substantially none of the aluminum-containing composition used to coat the internal passages contacts the external surface to deposit thereon. The inventors have found that deposition of the aluminum on the external surface, followed by deposition of a MCrAlX-type protective layer, may lead to reduced performance of the airfoil. The present approach also provides apparatus which permits the coating of only the internal surfaces with the aluminum and thence the aluminide.
An apparatus for coating a portion of a gas turbine airfoil having an external surface and an internal passage therethrough comprises a source of a flowable precursor coating material in contact with the internal passage of the airfoil, and a coating prevention structure overlying at least a portion of the external surface. The coating prevention structure prevents the contact of the flowable precursor coating material with the external surface of the airfoil.
In operation, a method for coating a portion of a gas turbine airfoil having an external surface and an internal passage therethrough comprises providing a source of a flowable precursor coating material in contact with the internal passage of the airfoil, providing a coating prevention structure overlying at least a portion of the external surface, and flowing the flowable precursor coating material from the source of the flowable precursor coating material and through the internal passage of the airfoil. The coating prevention structure substantially prevents contact of the flowable precursor coating material with the external surface of the airfoil.
The precursor coating material is preferably an aluminum-containing compound, optionally modified with the addition of a modifying element such as hafnium, zirconium, yttrium, silicon, titanium, tantalum, cobalt, chromium, platinum, and palladium, or combinations thereof. The source of the precursor coating material is preferably an aluminum halide gas, optionally mixed with source gases of the modifying elements.
The coating prevention structure substantially prevents contact of the precursor coating material with the external surface of the airfoil, both before the precursor coating material enters the internal passage and after it leaves the internal passage. The result is that very little, if any, of the precursor coating material contacts the external surface, so that little, if any, of the coating material that coats the internal passages is deposited on the external surface. In some cases, it may be desirable to coat the internal passages and a selected portion of the external surface with the aluminide coating, and not coat the remainder of the external surface 38 with the aluminide coating. The present invention provides an approach for achieving that type of coating.
Various types of external coating prevention structures may be used, singly or in combination. In one embodiment, the internal passage includes a precursor inlet end and a precursor outlet end, and the coating prevention structure comprises a housing that isolates the external surface from the precursor inlet end and the precursor outlet end. Where the airfoil has a platform, the housing may be used to isolate the lower side of the platform from the external surface. In other embodiments, the coating prevention structure may comprise a plurality of reactive particles that react with the precursor coating material, a solid mask, a slurry mask, a putty mask, and/or a flowing maskant gas. These various masking techniques may be used in combination, for example, a housing filled with the reactive particles adjacent to the coating surface.
The present approach produces a protective coating on the surfaces of the internal passages of the coating material, which may then be further processed as by interdiffusion. Substantially none of the coating material is deposited on the external surface of the airfoil section. Different protective layers may be deposited on the external surface, without interference from the protective coating on the internal surfaces.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings, which illustrate, by way of example, the principles of the invention. The scope of the invention is not, however, limited to this preferred embodiment.